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针对不同补燃室结构参数对固体火箭超燃冲压发动机补燃室掺混燃烧性能的影响进行研究,分析各级燃烧室的长度与扩张角度对补燃室性能的影响。采用基于密度的二阶迎风格式对补燃室掺混燃烧进行模拟,湍流模型和燃烧模型分别采用SST k-ω模型和涡团耗散模型。结果表明,提高燃烧效率与降低总压损失是相互矛盾的;燃烧效率随燃烧室长度的增加而增大,随燃烧室扩张角度的增加而减小;总压恢复系数随燃烧室长度的增加而减小,随燃烧室扩张角度的增加而增大;一级燃烧室的结构参数对燃烧效率与总压恢复系数的影响最大。当补燃室的总长与出口面积一定时,以发动机的总体性能参数作为补燃室构型的优化目标,对一、二级燃烧室长度与一、三级燃烧室扩张角度进行优化。
The effects of different combustion chamber structure parameters on the mixing combustion performance of the solid rocket scramjet were studied. The effects of combustion chamber length and expansion angle on the combustion chamber performance were analyzed. The second-order density-based upwind scheme was used to simulate the combustion in the afterburner. The turbulence model and the combustion model were respectively SST k-ω model and vortex dissipation model. The results show that it is contradictory to improve the combustion efficiency and reduce the total pressure loss. The combustion efficiency increases with the length of the combustion chamber and decreases with the expansion angle of the combustion chamber. The total pressure recovery coefficient increases with the length of the combustion chamber Decreases, and increases with the expansion angle of the combustion chamber. The structural parameters of the primary combustion chamber have the greatest impact on the combustion efficiency and the total pressure recovery coefficient. When the total length and the exit area of the afterburning chamber are fixed, the overall performance parameters of the engine are taken as the optimization objectives of the afterburning chamber configuration. The length of the first and second stage combustors and the expansion angles of the first and the third stage combustors are optimized.